Efficiency is a primary concern in the design of any gas turbine engine. Historically, one of the principle techniques for increasing efficiency has been to increase the gas path temperatures within the engine. Using internally cooled components made from high temperature capacity alloys has accommodated the increased temperatures. Turbine stator vanes and blades, for example, are typically cooled using compressor air worked to a higher pressure, but still at a lower temperature than that of the core gas flow passing by the blade or the vane. It will be understood that compressor bleed air for such cooling will be unavailable to support combustion in the combustor. The higher pressure provides the energy necessary to push the air through the component. A significant percentage of the work imparted to the air bled from the compressor, however, is lost during the cooling process. The lost work does not add to the thrust of the engine and negatively effects the overall efficiency of the engine. A person of skill in the art will recognize therefore, that there is a tension between the efficiency gained from higher core gas path temperatures and the concomitant need to cool turbine components and the efficiency lost from bleeding air to perform that cooling. There is, accordingly, great value in maximizing the cooling efficiency of whatever cooling air is used.
Thus, to minimize any sacrifice in engine performance due to unavailability of cooling airflow to support combustion, any scheme for cooling blades and vanes must optimize the utilization of compressor bleed cooling air. Airfoil cooling is accomplished by external film cooling, internal air impingement and forced convection either separately or a combination of all cooling methods.
In forced convection cooling, compressor bleed air flows through the internal cavities of the blades and vanes, continuously removing heat therefrom. Compressor bleed air enters the cavities 38 through one or more inlets which discharges into the internal cavities.
Film cooling has been shown to be very effective but requires a great deal of fluid flow to be bled off the compressor for cooling. Further, film cooling is actively controlled in a complex and expensive manner. Also, the fabrication and machining of an airfoil with film cooling holes adds a degree of complexity that is costly. It will also be appreciated that once the cooling air exits the internal cavity of the airfoil and mixes with the hot gases, a severe performance penalty is incurred due to the mixing process and the different temperature levels of the mixing flows. Thus, film cooling requires a greater amount of cooling air with the possibility of inadequate cooling of the outer surfaces of the airfoil.
Prior art coolable airfoils typically include a plurality of internal cavities (cooling circuit), which are supplied with cooling air. The cooling air passes through the wall of the airfoil (or the platform) and transfers thermal energy away from the airfoil in the process. Typically in the prior art, blade tip film cooling holes provide external film cooling issued on the blade tip pressure side in the radial and axial directions. Some designs use as many film holes as possible, in the limited space available, in an effort to flood the pressure side tip region with coolant. It is desired that this film cooling then carry over onto the outer tip surface to provide cooling there and also over the suction side surfaces of tip. Film holes are oriented in the radially outward direction because the prevailing mainstream gas flows tend to migrate in this manner in the tip region. In practice, it is still very difficult and very inconsistent to cool the blade tip in this manner due to the very complex nature of the cooling flow as it mixes with very dynamic hot gases of the mainstream flow.
In some prior art arrangements, cooling flow exits the film holes and is swept by the hot combustion gases towards the trailing edge of the airfoil and away from tip cap. Typically, this results in a mixed effect, where some of the cooling air is caught up and mixed with the hot gases and some goes onto tip cap and some goes axially along the airfoil to trailing edge. This results in inadequate cooling of tip cap and eventual temperature inflicted degradation of tip cap.
Turbine engine blade designers and engineers are constantly striving to develop more efficient ways of cooling the tips of the turbine blades to prolong turbine blade life and reduce engine operating cost. Cooling air used to accomplish this is expensive in terms of overall fuel consumption. Thus, more effective and efficient use of available cooling air in carrying out cooling of turbine blade tips is desirable not only to prolong turbine blade life but also to improve the efficiency of the engine as well, thereby again lowering engine operating cost. Consequently, there is a continuing need for a cooling design that will make more effective and efficient use of available cooling air.